Gas turbine engines are provided with rotor blades. The blades may be part of a rotor or a stator. The rotor blades may be provided in adjacent and distinct pairs that are radially offset from each other such that a front blade is radially forward of a rear blade. During operation of the gas turbine engine, compression of the flow is accomplished by turning the flow by the rotor blades. For a fixed rotational rotor speed the rotor blade loading is changing with mass flow rate such that it is higher at lower mass flow rate and lower at higher mass flow rate. Due to adverse pressure gradient the rotor blades may experience separation on suction side leading to stall limiting lower mass flow rate boundary of the rotor blade operational range. At last stages of high pressure compressor there is a leakage flow through clearances between rotating and stationary parts leading to degradation of performance. Reduced performance and operational ranges may result from the tip clearances being larger as compared to the rotor blade span than in the front stages due to smaller span size of rear stage rotor blades.
Accordingly, it is desirable to provide a blade having an improved solidity at the tip and improved gas turbine engine performance and operational ranges.